Fluid propulsion devices achieve thrust by imparting momentum to a fluid called the propellant. An air-breathing engine, as the name implies, uses the atmosphere for most of its propellant. The gas turbine produces high-temperature gas which may be used either to generate power for a propeller, fan, generator or other mechanical apparatus or to develop thrust directly by the expansion and acceleration of the hot gas in a nozzle. In any case, an air breathing engine continuously draws air from the atmosphere, compresses it, adds energy in the form of heat, and then expands it in order to convert the added energy to shaft work or jet kinetic energy. Thus, in addition to acting as propellant, the air acts as the working fluid in a thermodynamic process in which a portion of the energy is made available for propulsive purposes or work.
Typical gas turbine engines include at least two air streams. All air utilized by the engine initially passes through a fan prior to being split into the two or more air streams. The inner air stream is referred to as core air and passes into the compressor portion of the engine where it is compressed. The core air then is fed to the combustor portion of the engine where it is mixed with fuel and the fuel is combusted. The combustion gases are then expanded through the turbine portion of the engine that extracts energy from the hot combustion gases. The extracted energy is used to power the compressor, the fan and other accessory systems. The remaining hot gases then flow into the exhaust portion of the engine and may be used to produce thrust for the aircraft.
The outer air stream (also known as the bypass air stream) bypasses the engine core and is pressurized by the fan. Typically, no other work is done on the outer air stream as it flows axially in the engine but outside the core. Downstream of the turbine, the bypass air stream is used to cool engine hardware in the exhaust system. When additional thrust is required (or demanded), some of the bypass air stream may be redirected to an augmenter (afterburner) where it is mixed with the core air stream and fuel to provide the additional thrust to the aircraft. Referring to the drawings wherein similar reference numerals denote like elements throughout the various views, FIG. 1 shows a general orientation of a turbofan engine in a cut away view. In the turbofan engine shown, the flow of the air is generally axial. The position of components along the engine axis may be described using the terms “upstream” and “downstream” relative to the direction of airflow. Upstream refers to positions toward the engine inlet and downstream refers to positions toward the engine exhaust. Additionally, outward/outboard and inward/inboard refer to a lateral or radial direction perpendicular to the axial direction. For example, in the gas turbine engine of an aircraft, the bypass duct is outboard of the core duct. Typically the ducts are generally circular and co-axial.
The ambient inlet airflow 12 enters inlet fan duct 14 of turbofan engine 10, through the guide vanes 15, passes by fan spinner 16 and through fan rotor (fan blade) 42. The airflow 12 is split into primary (core) flow stream 28 and bypass flow stream 30 by upstream splitter 24 and downstream splitter 25. The bypass flow stream 30 along with the core/primary flow stream 28 is shown, the bypass stream 30 being outboard of the core stream 28. The inward portion of the bypass stream 30 and the outward portion of the core streams are partially defined by splitters 24 and 25 upstream of the compressor 26. The fan 42 has a plurality of fan blades.
As shown in FIG. 1 the fan blade 42 shown is rotating about the engine axis into the page, therefor the low pressure side of the blade 42 is shown, the high pressure side being on the opposite side of the blade. The high pressure and low pressure sides may also be known as the leading and trailing edges of the fan blade 42. The core flow stream 28 flows through compressor 26. The compressed air typically passes through an outlet guide vane to reduce or eliminate swirling motion or turbulence, a diffuser where air spreads out, and a compressor manifold to distribute the air in a smooth flow. The core flow stream 28 is then mixed with fuel in combustion chamber 35 and the mixture is ignited and burned. The resultant combustion products flow through turbines 38 that extract energy from the combustion gases to turn fan rotor 42, compressor 26 and provide any shaft work by way of turbine shaft. The gases, passing exhaust cone, expand through an exhaust nozzle 43 to produce thrust. The core flow stream 28 leaves the engine at a higher velocity than when it entered. The bypass flow stream 30 flows through fan rotor 42, flows by bypass duct outer wall 27 (an annular duct concentric with the core engine), flows through fan discharge outlet and is expanded through an exhaust nozzle to produce additional thrust. Turbofan engine 10 has a generally longitudinally extending centerline represented by engine axis 46.
The fan 42 generates a significant portion of the propulsive force for the turbofan engine 10. The fan 42 is large, composed of many attached fan blades and may rotate at high speeds to produce thrust. While the likelihood of a failure of one or more blades during operation is low, the consequences of a part of or a whole fan blade being thrown from a fan rotor during flight could led to catastrophic consequences for vital plane systems and passengers. To lessen the risks presented by a lost fan blade, turbine engines may use a containment system to capture any loose blades and absorb the energy from their impact. Such containment systems are particularly important in applications which use a composite containment case because composite material may be much more easily cut by a loose fan blade than a metal containment case.
A detailed illustration of a containment assembly of a turbine engine is shown in FIG. 2. The assembly 200 comprises a casing 202, a fan track liner 204, rotor 290, centerline axis 246, engine interface 208 and fan 242 (shown as a single fan blade). The fan blades 242 are attached to rotor 290 which spins about the centerline axis of the engine 246. The fan 242 spins into the page, as shown by arrow 216, compressing the incoming airstream 212.
Casing 202 provides support for the fan track liner 204 and may be connected to engine interface 208 via bolts 206 or other connection means. The casing 202 may be made of metal, composite, or a combination of the two materials. The casing 202 may be referred to as a fan-containment case. Composite materials are advantageous over metal casings in that they can support normal operating engine loads at greatly reduced weights. However, composite materials may be more susceptible to being damaged by, for example, a thrown fan blade than a metal containment case.
Due to this susceptibility to damage to casing 202, a fan track liner 204 is attached to the inner (inboard) radial wall of the casing 202 to absorb and redistribute the impact energy of a thrown blade, thereby avoiding localization of energy which would otherwise have been imparted onto the casing 202. A typical energy-absorbing fan track liner comprises a multilayer honeycomb and facesheet design which may be designed to fail in a staged sequence in order to protect the casing 202. The fan tracker liner 204 may be comprised of a plurality of sub-assemblies 318, for example, between five and eight, arranged radially outward from and circumferentially around the fan 242 as seen in FIG. 3. The liner 204 may extend axially forward and aft of furthest upstream and downstream part, respectively, of the fan blade to accommodate any axial movement of a blade after becoming loose from the fan. The fan track liner 204 is often bonded to the inner radial wall of the casing 202 such that the subassemblies 318 form a joint at which there is a discontinuity of energy-absorbing layers.
An axial cross-sectional view of a turbine engine 300 is shown in FIG. 3. This cross-section may be taken from a turbine engine such as that shown in FIG. 1 at section A-A. The engine 300 includes a casing 302, a fan track liner 304 and rotor 390. The casing 302 and fan track liner 304 may be commonly referred to as a containment assembly 322. The rotor spins in a counter-clockwise direction about the centerline axis 346 as shown by arrow 316. A fan is not shown because the cross-section is taken just forward of the fan. The inlet air flow direction is shown by arrow 312.
The casing 302 circumscribes the rotor 390 and is outboard of the rotor 390 and the fan track liner 304. The casing 302 may comprise materials as described above.
The fan track liner 304 circumscribes the rotor 390 and is inboard of the casing 302. The fan tracker liner 304 may comprise a plurality of subassemblies 318, also known as panels or crush panels, for example, five or eight subassemblies. The fan track liner is divided into subassemblies for manufacturing and engine assembly purposes. The subassemblies 318 may comprise the materials and construction as described above.
As can be seen in FIG. 3, the subassemblies 318 join one other at a series of joints 320 located circumferentially around the axis 346. These joints 320 are end-to-end butt joints with the joint 320 oriented in the radial direction. A close up of an end-to-end butt joint in a containment assembly 400 is shown in FIG. 4. The containment assembly 422 comprises casing 402 and fan track liner 404, liner 404 comprising a plurality of subassemblies 418. The subassemblies 418 meet at an end-to-end butt joint 420. The fan 442 rotates counterclockwise around the centerline axis (not shown) as indicated by direction arrow 416. Subassemblies 418 may be of a multilayer design comprising facesheet 432 and an energy-absorbing layer 434, which may be of a metallic honeycomb design.
The end-to-end butt joint method of joining the subassemblies 318/418 provides a discontinuity in the facesheet and/or honeycomb material running circumferentially around the engine, leading to a sharp drop in stiffness at the joint location that can permit the blade to penetrate to deeper layers at the joint. This discontinuity severely impacts the transfer of blade load across the joint surface, resulting in blade impact behavior that differs at the joint 320/420 as opposed to at the center of the panels 318/418. Other designs use a perpendicular or adjacent ninety degree upturn joint which may lead to an overly stiff joint that may concentrate an impact load on a small area of the barrel (or casing) in the vicinity of the joint, thereby potentially causing damage.
With reference to FIG. 2, the energy-absorbing effectiveness of the fan track liner 204 is highly correlated to the overall thickness of the liner 204. However, the maximum thickness of the liner is limited by overall plane design criteria to include engine nacelle size limitations, casing 202 load bearing requirements and the size of the engine equipment. Some designs create an “arched” casing 202 in which the casing inner wall proximate to the fan track liner section is pushed outwardly. While such a design allows for thicker fan track liners, these casing must be formed by fusing at least two separate components because the entire arched-design casing cannot be removed from the mandrel around which it is modeled without cutting the casing or destroying the mandrel. A casing formed by fusing two separate sections is inherently weaker than a casing formed from continuous plies.
As disclosed in some embodiments herein the current subject matter addresses these deficiencies by utilizing an angled upturn joint between adjoining subassemblies (or panels) of a fan track liner. The angled upturn joint allows smooth transition of a crushing load from a leading subassembly panel to the energy absorbing layers of a trailing subassembly panel to which the leading subassembly panel is joined. The angled upturn joint further mitigates damage to the casing to which the subassemblies are attached when compared to joining methods which create large discontinuities in stiffness at the joints. The use of an angled upturn joint also permits co-curing of the septum resin and the energy-absorbing layer adhesive, thereby enabling a certain amount of consolidation of the inner septum at the angled upturn joint.
The disclosed subject matter in accordance with some embodiments also addresses the limitations on fan track liner thickness and two-part casings by embedding a layer of the fan track liner within the layup of the casing. This creates a containment casing have an embedded containment core. The load carrying fibers of the containment casing having an embedded core may be pushed to the outboard side of the casing. The inner fibers, or plies, effectively serve as a load transfer septum between the energy-absorbing layers of the fan tracker liner bonded to the inner radial surface of the containment casing having an the embedded energy absorbing layer, and may replace the bespoke intermediate glass septum used in other applications. This may maximize the thickness of the impact energy absorbing fan track liner, and may take advantage of additive manufacturing methods of composite component fabrication by including the layer of energy-absorbing honeycomb into the composite layup. Some embodiments may include either aramid fiber or metallic honeycomb.
In accordance with some embodiments of the present disclosure, a machine having a containment assembly is provided. The machine may have a rotor that is at least partially bounded by the containment assembly. The containment assembly may comprise an inner casing member, a containment liner, an outer casing member, and a containment core. The inner casing member may be formed from a composite material and have a radially inner surface facing the rotor. The containment liner may be bonded to the radially inner surface of the inner casing member. The outer casing member may be formed from composite material and be positioned radially outward of the inner casing member. The containment core may be embedded between the inner and outer casing members.
In accordance with some embodiments of the present disclosure, a method of making a containment assembly is provided. The containment assembly may radially bound a rotor of a machine. The method may comprise the steps of forming a casing on a mandrel from a composite material, curing the composite material, removing the casing from the mandrel, and bonding a containment liner to the radially inner surface of the casing. The step of forming the casing may comprise forming an inner casing member from a composite material on the surface of the mandrel, positioning a containment core on the outer surface of the inner casing member, and forming an outer casing member from a composite material so that the containment core is embedded between the inner and outer casing members.
In accordance with some embodiments of the present disclosure, a containment assembly is provided. The containment assembly may bound at least in part a turbine engine having a rotor. The containment assembly may comprise a casing, a fan track liner, an outer radial surface, and a containment core. The casing is formed from a composite material and may have a radially inner surface facing the rotor. The fan track liner may be bonded to the radially inner surface of said inner casing member. The outer radial surface may be formed from composite material and be positioned radially outward of said inner surface. The containment core may be embedded between the inner and outer surfaces.
These and many other advantages of the present subject matter will be readily apparent to one skilled in the art to which the disclosure pertains from a perusal of the claims, the appended drawings, and the following detailed description of preferred embodiments.